"The government of the United States of America has rights in this invention pursuant to Contract No. DE-AC02-92CE40960 awarded by the United States Department of Energy".
In the operation of a gas turbine engine, air at atmospheric pressure is initially compressed by a compressor, and the resulting compressed air is delivered to a combustion stage. In the combustion stage, heat is added to the compressed air leaving the compressor by mixing fuel with the compressed air and by burning the fuel/air mixture. The gas flow resulting from combustion of the fuel/air mixture in the combustion stage expands through a turbine, and some of the energy of the gas flow is used to drive a turbine in order to produce mechanical power.
One form of turbine is an axial turbine having one or more stages, wherein each stage employs one row of stationary nozzle guide vanes and one row of moving blades. The row of moving blades is mounted on a turbine disk. The nozzle guide vanes are aerodynamically designed to direct incoming gas from the combustion stage onto the turbine blades to thereby aerodynamically transfer kinetic energy to the blades.
The combustion gases entering the turbine typically have a gas entry temperature in the range of 850.degree. C. to at least 1250.degree. C. Since the efficiency and work output of the turbine engine are related to the gas entry temperature of the incoming combustion gases, there is a trend in gas turbine engine technology to increase the gas entry temperature. Seals are normally provided on the disk of a high temperature turbine wheel either near the hub of the turbine or near the rim of the disk of the turbine wheel in order to confine these hot gases of the gas turbine engine to a predetermined area of the turbine. If the seals are provided on the disk of the turbine wheel near the hub of the turbine, the hot gases contact a substantially greater surface area of the disk and turbine than if the seals are provided near the rim of the disk of the turbine wheel. Consequently, providing the seals near the hub of the turbine requires a seal support diaphragm and substantially more cooling than is required by providing effective seals near the rim of the disk of the turbine wheel. Therefore, it is advantageous to mount the seals near the rim of the disk of a turbine wheel.
One typical form of a rim seal is a labyrinth rim seal which is mounted on the disk of the turbine wheel near the rim of the disk and which cooperates with a metal stator ring fixed to the housing of the turbine. The labyrinth rim seal, together with the metal stator ring, isolate the hot gases of the gas turbine engine from all or most of the disk area.
The clearance between the labyrinth rim seal affixed to the disk of the turbine wheel and the metal stator ring supported by the housing of the turbine must be large enough to accommodate the different rates of thermal expansion and thermal shrinkage of the metal stator ring against the labyrinth rim seal. These different rates of thermal expansion and thermal shrinkage result, at least in part, from the relative materials, the relative thermal expansion coefficients, and the different masses of the metal stator ring, the labyrinth rim seal, and the structures on which the metal stator ring and the labyrinth rim seal are supported.
That is, a gas turbine engine is manufactured with a room temperature clearance between the metal stator ring and the labyrinth rim seal. During initial operation of the gas turbine engine, the metal stator ring usually expands more rapidly than the labyrinth rim seal primarily because the metal stator ring and its supporting structure has a typically lower mass than does the labyrinth rim seal and the disk to which the labyrinth rim seal is attached. Therefore, the clearance between the metal stator ring and the labyrinth rim seal increases during engine start-up. Then, as the larger mass of the labyrinth rim seal and the disk to which the labyrinth rim seal is attached continues to heat up, the clearance between the labyrinth rim seal and the metal stator ring reduces to a steady state condition.
During shut down of the turbine engine, the metal stator ring shrinks more rapidly than the labyrinth rim seal primarily because the metal stator ring and its supporting structure has a lower mass than does the labyrinth rim seal and the disk to which the labyrinth rim seal is attached. Therefore, the clearance between the metal stator ring and the labyrinth rim seal shrinks below the steady state clearance. Then, after the gas turbine engine is completely cooled, the clearance between the metal stator ring and the labyrinth rim seal resumes the size of the room temperature clearance.
The steady state clearance between the metal stator ring and the labyrinth rim seal must be sufficiently large to accommodate the different expansion and shrink rates between the lower mass metal stator ring and the higher mass labyrinth rim seal as the gas turbine engine is brought up to, and is shut down from, its steady state operation. However, this large clearance impairs the quality of the seal between the metal stator ring and the labyrinth rim seal during steady state operation of the gas turbine engine. This large clearance results in large quantities of hot gas to flow from the gas path of the gas turbine engine into the cavity surrounding the disk of the turbine wheel. Furthermore, the different expansion rates between the lower mass metal stator ring and the higher mass labyrinth rim seal during initial operation of the gas turbine engine allows even more hot gas from the gas path of the gas turbine engine to flow into the cavity surrounding the disk of the turbine wheel. Therefore, the turbine disk requires a large amount of cooling, and corresponding performance penalties result.
The present invention is directed to overcome one or more of the problems as set forth above.